1. Field of the Invention
The invention concerns a device and a method for calibrating rate gyro drift for a satellite in equatorial or near-equatorial orbit (typically an orbit at less than approximately 10.degree. to the plane of the equator) having at least one Sun detector, one Earth or Pole Star detector and rate gyros (of the integrating type or otherwise) and stabilized with respect to three axes by a set of thrusters.
2. Description of the Prior Art
The three axes with respect to which the attitude of a satellite is stabilized are respectively a yaw axis Z, a pitch axis Y and a roll axis X completing the direct orthonomic frame of reference. Stabilizing the attitude of the satellite entails keeping this frame of reference parallel to (or even coincident with) a theoretical (or reference) frame of reference which, in the case of a circular orbit, for example, is an axis Zr directed towards the Earth, an axis Yr perpendicular to the plane of the orbit and an axis Xr in the same direction as the orbital speed of the satellite. Other types of reference frame of reference are conventionally used to suit the type of orbit and the type of sensors with which the satellite is provided. The theoretical frame of reference Xo, Yo, Zo on which the X, Y, Z frame of reference of the satellite is to be superimposed is called the instantaneous local orbital frame of reference (it is dependent on the point in the orbit at which the satellite is located at the time).
Any telecommunication or observation satellite (these are typical satellite missions) stabilized with respect to three axes in the transfer orbit and in its final orbit has at least one solar sensor (one-axis or two-axis), one two-axis Earth sensor or one two-axis Pole Star sensor for detecting attitude errors (in roll and in pitch in the case of the Earth sensor, in roll and in yaw in the case of the Pole Star sensor, and in yaw, roll and pitch in the case of the solar sensor), and rate gyros for measuring the angular speed around each of the three axes.
Command torque for changing the attitude and/or the angular speeds of the satellite are usually generated by actuators such as momentum or reaction wheels or thrusters. This type of satellite carries at least one apogee motor (in practice a thruster) to travel to a nominal service orbit from the transfer orbit and a set of thrusters for thereafter (i.e. during the operational phase) maintaining the satellite in orbit and in an appropriate attitude.
The set of sensors and actuators associated with one or more electronic systems for processing the measured values and computing command torques constitutes what is usually referred to as the attitude and orbit control system (AOCS) of the satellite.
It is essential to be able to estimate at any time during the service life of the satellite the drift affecting the rate gyros, or at least the constant part (bias) of such drift. For a satellite carrying an Earth sensor, in particular, when the satellite is in the transfer orbit and during the orbit change phase--apogee thrust mode--the terrestrial sensor may no longer be able to see the Earth and it may be necessary to maintain the attitude of the satellite using only measurements by the other sensors, in particular the solar sensor(s) and the rate gyros (integrating or otherwise).
Likewise, during station-keeping maneuvers in the nominal orbit it is preferable to be able to measure the angular speeds of the satellite accurately in order to damp quickly any attitude disturbances caused by misalignment of the station-keeping thrusters, as such misalignments are difficult to avoid in practice.
In all these cases, attitude control stability is dependent on measurement of the angular speed by the rate gyros. In practice these measured values are often subject to a bias because of rate gyro drift, the order of magnitude of the bias being typically a few degrees per hour, and this can substantially degrade control performance.
Further, the consequences of such drift are even more important if the measured values are integrated (substituting an optical sensor for position measurements) without compensating the bias for a certain time: the result is a large error in the estimated attitude angles and consequently degraded pointing of the satellite.
For example, for a typical orbit change maneuver from the transfer orbit the satellite attitude must be maintained for approximately one hour using the rate gyros and the solar sensors. If rate gyro drift is not compensated, the resulting error in the orientation of the thrust can move the satellite to an intermediate orbit very different from that required (the error can be as much as several hundred kilometers for a drift of a few .degree./h).
To prevent this from happening rate gyros must in practice be calibrated at least once in the transfer orbit and possibly subsequently in the final orbit. The method and device of the present invention are concerned with this calibration.
Rate gyro calibration concepts for a three-axis stabilized satellite using solar and terrestrial sensors are known and described, for example in French patent 2,583,873 or European patent 209,429 covering a method and device for injecting a satellite into a geostationary orbit with stabilization about three axes and French patent 2,622,001 (or German patent 3,734,941 or U.S. Pat. No. 4,884,771) "Calibrating gyroscope of three-axis stabilized satellite".
In the first patent only drift of the roll and yaw rate gyros are calibrated by maintaining the yaw axis of the satellite parallel to itself. The second patent describes a similar procedure for estimating the drift of a rate gyro for any axis, the satellite being aligned with two successive reference positions.
Satellite attitude control systems using a star sensor in addition to solar and terrestrial sensors are known and described, for example, in French patent 2,522,614 covering an equatorial orbit satellite configuration with improved solar means and French patent 2,637,565 covering a three-axis active control system for a geostationary satellite. However, there is no provision in these two patents for calibration of rate gyro drift whether in the transfer orbit or on station.
The use of conventional filtering methods such as the Kalman filter is proposed in "On-Orbit Attitude Reference Alignment and Calibration", AAS paper No. 90-042, for conjointly estimating satellite attitude errors, star sensor alignment errors and scale factors, misalignments and constant drift (bias) of a three-axis rate gyro system. This document teaches storing star sensor measurements and rate gyro measurements during pointing maneuvers prescribed for the mission and processing them off-line (either on board or on the ground) in order to reconstitute the attitude of the satellite and estimate sensor misalignment and rate gyro misalignment, scale factor and drift.